The present application is related to Ser. No. 13/479,710 and Ser. No. 13/479,683, filed concurrently herewith, which are fully incorporated by reference herein and made a part hereof.
The present application relates generally to apparatus, methods and/or systems for cooling the tips of gas turbine rotor blades. More specifically, but not by way of limitation, the present application relates to apparatus, methods and/or systems related to microchannel design and implementation in turbine blade tips.
In a gas turbine engine, it is well known that air is pressurized in a compressor and used to combust a fuel in a combustor to generate a flow of hot combustion gases, whereupon such gases flow downstream through one or more turbines so that energy can be extracted therefrom. In accordance with such a turbine, generally, rows of circumferentially spaced rotor blades extend radially outwardly from a supporting rotor disk. Each blade typically includes a dovetail that permits assembly and disassembly of the blade in a corresponding dovetail slot in the rotor disk, as well as an airfoil that extends radially outwardly from the dovetail.
The airfoil has a generally concave pressure side and generally convex suction side extending axially between corresponding leading and trailing edges and radially between a root and a tip. It will be understood that the blade tip is spaced closely to a radially outer turbine shroud for minimizing leakage therebetween of the combustion gases flowing downstream between the turbine blades. Maximum efficiency of the engine is obtained by minimizing the tip clearance or gap such that leakage is prevented, but this strategy is limited somewhat by the different thermal and mechanical expansion and contraction rates between the rotor blades and the turbine shroud and the motivation to avoid an undesirable scenario of having excessive tip rub against the shroud during operation.
In addition, because turbine blades are bathed in hot combustion gases, effective cooling is required for ensuring a useful part life. Typically, the blade airfoils are hollow and disposed in flow communication with the compressor so that a portion of pressurized air bled therefrom is received for use in cooling the airfoils. Airfoil cooling is quite sophisticated and may be employed using various forms of internal cooling channels and features, as well as cooling holes through the outer walls of the airfoil for discharging the cooling air. Nevertheless, airfoil tips are particularly difficult to cool since they are located directly adjacent to the turbine shroud and are heated by the hot combustion gases that flow through the tip gap. Accordingly, a portion of the air channeled inside the airfoil of the blade is typically discharged through the tip for the cooling thereof.
It will be appreciated that conventional blade tip design includes several different geometries and configurations that are meant to prevent leakage and increase cooling effectiveness. Exemplary patents include: U.S. Pat. No. 5,261,789 to Butts et al.; U.S. Pat. No. 6,179,556 to Bunker; U.S. Pat. No. 6,190,129 to Mayer et al.; and, U.S. Pat. No. 6,059,530 to Lee. Conventional blade tip designs, however, all have certain shortcomings, including a general failure to adequately reduce leakage and/or allow for efficient tip cooling that minimizes the use of efficiency-robbing compressor bypass air. In addition, as discussed in more detail below, conventional blade tip design, particularly those having a “squealer tip” design, have failed to take advantage of or effectively integrate the benefits of microchannel cooling. As a result, an improved turbine blade tip design that increases the overall effectiveness of the coolant directed to this region would be in great demand.